The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which powers the compressor.
And, additional energy is extracted in a low pressure turbine (LPT) which drives an upstream fan in an aircraft turbofan aircraft engine application, or drives an external drive shaft in marine and industrial applications.
The modern combustor is annular and includes radially outer and inner combustion liners extending downstream from a forward dome to define an annular combustion zone. A row of fuel injectors and cooperating air swirl cups are mounted in the dome for discharging air atomized fuel jets that are suitably ignited for generating the combustion gases.
The fuel injectors are spaced circumferentially apart from each other typically in a uniform distribution, and correspondingly effect relatively hot streaks of combustion gases which flow downstream to the annular combustor outlet.
The maximum combustion gas temperature is found along the center of each hot streak, and the combustion gas temperature correspondingly decreases radially outwardly from the centerline of each hot streak, which is both radially between the outer and inner combustor liners, as well as circumferentially around the combustor between the circumferentially spaced apart hot streaks.
The resulting temperature pattern of the combustion gases at the annular combustor outlet varies both radially between the outer and inner liners, and circumferentially between the hot streaks, with the lower temperature gases between the hot streaks typically being referred to as cold streaks. The differential temperature between the hot and cold streaks may be several hundreds of degrees and affects performance and operation of the downstream turbine components.
More specifically, the combustion gases discharged from the combustor outlet are first received by the first stage HPT turbine nozzle which guides the gases to the following first stage row of turbine rotor blades mounted on the perimeter of a supporting rotor disk. The turbine nozzle includes a row of hollow nozzle vanes mounted radially between corresponding outer and inner bands.
The nozzle is typically segmented circumferentially in a common configuration of nozzle doublets having two vanes integrally mounted in corresponding outer and inner band segments.
The annular nozzle is therefore circumferentially divided by axial splitlines at corresponding endfaces of the outer and inner bands of the nozzle doublets. And, the endfaces typically include slots for mounting spline seals therein for maintaining the circumferential continuity of the turbine nozzle and sealing internal cooling air loss therefrom.
The number of nozzle vanes in the complete row is substantially greater than the number of fuel injectors in the combustor and is commonly not an integer multiple thereof. Accordingly, in the assembly of the combustor relative to the turbine nozzle, the fuel injectors vary in relative circumferential position with the leading edges of the row of nozzle vanes.
The hot streaks generated from the fuel injectors during operation are therefore circumferentially aligned or clocked differently or randomly from vane to vane, and therefore subject the vanes to different heat loads during operation. The hot streaks bathe the nozzle vanes in maximum temperature combustion gases, whereas the circumferentially intervening cold streaks bathe the vanes in relatively cooler combustion gases.
Accordingly, the turbine nozzle is commonly designed with circumferential uniformity having substantially identical nozzle vanes and band segments, in the typical doublet configuration for example. An even number of nozzle vanes is therefore found in the doublet nozzle configuration with two identical vanes in each doublet.
The nozzle vanes have the typical crescent profile with generally concave pressure sides and generally convex suction sides extending axially in chord between opposite leading and trailing edges. The vanes in each doublet define an inboard flow passage therebetween, with the vanes between doublets defining outboard flow passages which include the respective axial splitlines.
The inboard and outboard nozzle passages converge in the downstream direction to a minimum flow area typically defined at the trailing edge of one vane normal to the suction side of the adjacent vane.
The combustion gases are typically discharged at an oblique circumferential swirl angle into the downstream row of turbine rotor blades which rotate the supporting rotor disk in the direction of the blade suction sides relative to the blade pressure sides.
Each nozzle doublet therefore includes a lead vane over which the turbine blades first pass, and a trail vane over which the turbine blades secondly pass during rotation.
The cold and hot streaks from the combustor are channeled axially through the flow passages of the turbine nozzle and therefore similarly bathe the turbine rotor blades in the alternating hot and cold streaks which also affects their performance during operation.
Surrounding the turbine blades is an annular turbine shroud which confines the combustion gases, including the hot and cold streaks. And, the shroud is also segmented circumferentially with identical turbine shroud segments having corresponding hooks supported in a cooperating hanger suspended from a surrounding casing or shroud support.
Accordingly, the nozzle stator vanes, turbine rotor blades, and their shrouds are typically identical in each row thereof and typically include identical cooling circuits therein for their different environments. The vanes, blades, and shrouds use a portion of pressurized air bled from the compressor for cooling thereof and achieving the desired useful life of the engine during operation.
Since the air bled from the compressor is not used in the combustor, the overall efficiency of the engine is decreased. The amount of cooling air bled from the compressor should therefore be minimized for maximizing engine efficiency.
However, the vanes, blades, and shrouds must be designed in conventional practice for identical cooling thereof in each row for protecting the airfoils from the maximum temperatures and heat loads from the hot streaks produced by the combustor notwithstanding the significantly lower temperature of the cold streaks alternating with the hot streaks during operation.
Accordingly, it is desired to provide an improved turbine which preferentially accommodates the hot and cold streaks in the combustion gases for improving performance of the gas turbine engine.